EPPLER 335 AIRFOIL (e335-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER 335 AIRFOIL (e335-il) Reynolds number: 100,000 Max Cl/Cd: 44.18 at α=8.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e335-il-100000-n5.txt Download as CSV file: xf-e335-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 335 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5093 0.09241 0.08777 -0.0105 1.0000 0.0231
-9.500 -0.5086 0.08878 0.08418 -0.0119 1.0000 0.0226
-9.250 -0.5156 0.08389 0.07931 -0.0148 1.0000 0.0220
-9.000 -0.5271 0.07913 0.07454 -0.0166 1.0000 0.0216
-8.500 -0.5726 0.06814 0.06326 -0.0149 1.0000 0.0189
-8.250 -0.5791 0.06539 0.06047 -0.0126 1.0000 0.0188
-8.000 -0.5832 0.06240 0.05740 -0.0104 1.0000 0.0186
-7.750 -0.5867 0.05916 0.05403 -0.0079 1.0000 0.0184
-7.500 -0.5876 0.05599 0.05071 -0.0054 1.0000 0.0182
-7.250 -0.5863 0.05288 0.04742 -0.0027 1.0000 0.0179
-7.000 -0.5852 0.04938 0.04367 0.0005 1.0000 0.0180
-6.750 -0.5814 0.04615 0.04017 0.0037 1.0000 0.0182
-6.500 -0.5770 0.04283 0.03647 0.0073 1.0000 0.0186
-6.250 -0.5617 0.03922 0.03235 0.0090 0.9572 0.0189
-6.000 -0.5283 0.03635 0.02912 0.0069 0.8966 0.0197
-5.750 -0.5015 0.03365 0.02591 0.0071 0.8538 0.0201
-5.500 -0.4788 0.03121 0.02292 0.0087 0.8198 0.0202
-5.250 -0.4554 0.02922 0.02046 0.0101 0.7911 0.0202
-5.000 -0.4299 0.02739 0.01822 0.0112 0.7669 0.0207
-4.750 -0.4020 0.02567 0.01613 0.0119 0.7445 0.0214
-4.500 -0.3727 0.02417 0.01432 0.0122 0.7242 0.0222
-4.250 -0.3439 0.02291 0.01281 0.0126 0.7058 0.0236
-4.000 -0.3186 0.02207 0.01190 0.0131 0.6885 0.0264
-3.750 -0.2932 0.02116 0.01082 0.0139 0.6722 0.0291
-3.500 -0.2705 0.02033 0.00992 0.0151 0.6575 0.0319
-3.250 -0.2484 0.01961 0.00910 0.0164 0.6432 0.0371
-3.000 -0.2258 0.01897 0.00839 0.0177 0.6296 0.0461
-2.750 -0.1168 0.01666 0.00951 0.0049 0.6112 0.7939
-2.500 -0.1068 0.01771 0.01036 0.0101 0.5996 0.8671
-2.250 -0.0493 0.01936 0.01161 0.0069 0.5848 0.9121
-2.000 0.0241 0.01996 0.01180 -0.0006 0.5694 0.9387
-1.750 0.0588 0.01991 0.01147 -0.0020 0.5584 0.9469
-1.500 0.0913 0.01974 0.01109 -0.0031 0.5474 0.9514
-1.250 0.1185 0.01969 0.01085 -0.0032 0.5373 0.9570
-1.000 0.1496 0.01954 0.01048 -0.0040 0.5284 0.9608
-0.750 0.1816 0.01937 0.01017 -0.0052 0.5183 0.9645
-0.500 0.2105 0.01929 0.00992 -0.0056 0.5100 0.9687
-0.250 0.2386 0.01923 0.00973 -0.0060 0.5013 0.9727
0.000 0.2707 0.01906 0.00943 -0.0071 0.4934 0.9757
0.250 0.3008 0.01896 0.00922 -0.0079 0.4854 0.9790
0.500 0.3291 0.01893 0.00908 -0.0083 0.4781 0.9825
0.750 0.3590 0.01884 0.00892 -0.0090 0.4706 0.9852
1.000 0.3895 0.01874 0.00870 -0.0099 0.4642 0.9878
1.250 0.4192 0.01868 0.00862 -0.0106 0.4568 0.9905
1.500 0.4479 0.01867 0.00852 -0.0111 0.4508 0.9932
1.750 0.4778 0.01862 0.00846 -0.0119 0.4439 0.9955
2.000 0.5074 0.01858 0.00838 -0.0126 0.4378 0.9977
2.250 0.5370 0.01858 0.00833 -0.0133 0.4321 1.0000
2.500 0.5603 0.01869 0.00848 -0.0128 0.4259 1.0000
2.750 0.5834 0.01881 0.00856 -0.0121 0.4208 1.0000
3.000 0.6066 0.01896 0.00870 -0.0115 0.4155 1.0000
3.250 0.6297 0.01911 0.00890 -0.0109 0.4097 1.0000
3.500 0.6528 0.01926 0.00904 -0.0102 0.4048 1.0000
3.750 0.6759 0.01945 0.00926 -0.0096 0.3998 1.0000
4.000 0.6989 0.01964 0.00952 -0.0090 0.3940 1.0000
4.250 0.7218 0.01982 0.00969 -0.0083 0.3892 1.0000
4.500 0.7447 0.02004 0.00999 -0.0077 0.3841 1.0000
4.750 0.7674 0.02027 0.01032 -0.0070 0.3785 1.0000
5.000 0.7901 0.02047 0.01052 -0.0063 0.3737 1.0000
5.250 0.8127 0.02074 0.01089 -0.0056 0.3684 1.0000
5.500 0.8349 0.02100 0.01127 -0.0049 0.3627 1.0000
5.750 0.8574 0.02121 0.01149 -0.0041 0.3579 1.0000
6.000 0.8793 0.02153 0.01196 -0.0034 0.3524 1.0000
6.250 0.9011 0.02182 0.01237 -0.0026 0.3465 1.0000
6.500 0.9233 0.02203 0.01259 -0.0018 0.3418 1.0000
6.750 0.9440 0.02242 0.01319 -0.0009 0.3352 1.0000
7.000 0.9653 0.02268 0.01354 0.0000 0.3293 1.0000
7.250 0.9863 0.02298 0.01393 0.0010 0.3237 1.0000
7.500 1.0062 0.02335 0.01449 0.0020 0.3167 1.0000
7.750 1.0272 0.02355 0.01472 0.0030 0.3112 1.0000
8.000 1.0457 0.02401 0.01543 0.0042 0.3033 1.0000
8.250 1.0656 0.02424 0.01572 0.0053 0.2969 1.0000
8.500 1.0831 0.02470 0.01641 0.0066 0.2888 1.0000
8.750 1.1018 0.02494 0.01672 0.0080 0.2817 1.0000
9.000 1.1178 0.02544 0.01745 0.0094 0.2728 1.0000
9.250 1.1346 0.02576 0.01785 0.0110 0.2650 1.0000
9.500 1.1491 0.02626 0.01856 0.0127 0.2555 1.0000
9.750 1.1627 0.02677 0.01923 0.0146 0.2461 1.0000
10.000 1.1755 0.02723 0.01975 0.0166 0.2371 1.0000
10.250 1.1853 0.02790 0.02061 0.0188 0.2265 1.0000
10.500 1.1931 0.02863 0.02148 0.0213 0.2162 1.0000
10.750 1.1981 0.02941 0.02234 0.0242 0.2062 1.0000
11.000 1.1993 0.03030 0.02331 0.0275 0.1965 1.0000
11.250 1.1961 0.03136 0.02449 0.0312 0.1870 1.0000
11.500 1.1858 0.03243 0.02560 0.0358 0.1799 1.0000
11.750 1.1674 0.03368 0.02694 0.0412 0.1738 1.0000
12.000 1.1488 0.03529 0.02856 0.0454 0.1688 1.0000
12.250 1.1310 0.03744 0.03080 0.0485 0.1627 1.0000
12.500 1.1137 0.03993 0.03333 0.0507 0.1569 1.0000
12.750 1.0977 0.04273 0.03617 0.0522 0.1513 1.0000
13.000 1.0814 0.04591 0.03941 0.0532 0.1452 1.0000
13.250 1.0664 0.04917 0.04266 0.0538 0.1396 1.0000
13.500 1.0512 0.05281 0.04639 0.0538 0.1336 1.0000
14.000 1.0239 0.06020 0.05384 0.0535 0.1222 1.0000
14.250 1.0115 0.06390 0.05753 0.0530 0.1168 1.0000
14.500 0.9998 0.06773 0.06141 0.0524 0.1112 1.0000
14.750 0.9889 0.07167 0.06540 0.0516 0.1058 1.0000
15.000 0.9804 0.07535 0.06906 0.0509 0.1007 1.0000
15.250 0.9708 0.07950 0.07331 0.0498 0.0956 1.0000
15.500 0.9644 0.08306 0.07683 0.0489 0.0907 1.0000
15.750 0.9552 0.08744 0.08132 0.0475 0.0859 1.0000
16.000 0.9479 0.09151 0.08544 0.0462 0.0814 1.0000
16.250 0.9429 0.09522 0.08914 0.0451 0.0771 1.0000
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